Inlet assembly for an aircraft aft fan

ABSTRACT

The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members, such as struts or strakes, mounted at predetermined locations around a circumference of the fan shaft of the fan at the inlet. The predetermined location(s) may be determined as a function of swirl distortion entering the inlet. As such, the structural member(s) are configured to reduce swirl distortion of the airflow entering the fan. In some embodiments, the inlet assembly may also include inlet guide vanes. In alternative embodiments, the inlet assembly may be absent of inlet guide vanes.

FIELD OF THE INVENTION

The present subject matter relates generally to aft engines for aircraftpropulsion systems, and more particularly to an inlet assembly for anaft fan of the aft engine that reduces swirl distortion of airflowentering the fan.

BACKGROUND OF THE INVENTION

A conventional commercial aircraft generally includes a fuselage, a pairof wings, and a propulsion system that provides thrust. The propulsionsystem typically includes at least two aircraft engines, such asturbofan jet engines. Each turbofan jet engine is mounted to arespective one of the wings of the aircraft, such as in a suspendedposition beneath the wing, separated from the wing and the fuselage.Such a configuration allows for the turbofan jet engines to interactwith separate, freestream airflows that are not impacted by the wingsand/or fuselage. This configuration can reduce an amount of turbulencewithin the air entering an inlet of each respective turbofan jet engine,which has a positive effect on a net propulsive thrust of the aircraft.

However, a drag on the aircraft including the turbofan jet engines alsohas an effect on the net propulsive thrust of the aircraft. A totalamount of drag on the aircraft, including skin friction, form, andinduced drag, is generally proportional to a difference between afreestream velocity of air approaching the aircraft and an averagevelocity of a wake downstream from the aircraft that is produced due tothe drag on the aircraft.

As such, systems have been proposed to counter the effects of dragand/or to improve an efficiency of the turbofan jet engines. Forexample, certain propulsion systems incorporate boundary layer ingestionsystems to route a portion of relatively slow moving air forming aboundary layer across, e.g., the fuselage and/or the wings, into theturbofan jet engines. Although this configuration can reduce drag byreenergizing the boundary layer airflow downstream from the aircraft,the relatively slow moving flow of air from the boundary layer enteringthe turbofan jet engine generally has a non-uniform or distortedvelocity profile. As a result, such turbofan jet engines can experiencean efficiency loss minimizing or negating any benefits of reduced dragon the aircraft.

In addition, some propulsion systems include an electrically-driven aftengine having an aft fan on the aircraft empennage to derive propulsivebenefit by ingesting fuselage boundary layers. During operation, theinlet of the aft fan can see a strong swirl distortion due to upwardflow from the bottom of the fuselage to the top. The swirl distortioncan be detrimental to fan operability and can cause aeromechanicaland/or operational issues.

Thus, an improved inlet assembly for an aft fan that addresses theaforementioned issue would be welcomed in the art. More particularly, aninlet assembly of the aft fan that reduces the swirl distortion would beespecially beneficial.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one aspect, the present disclosure is directed to a boundary layeringestion fan assembly for mounting to an aft end of a fuselage of anaircraft. The boundary layer ingestion fan assembly includes a fanrotatable about a central axis of the boundary layer ingestion fan. Morespecifically, the fan includes a plurality of fan blades rotatable abouta fan shaft that extends along the central axis. The fan assembly alsoincludes a nacelle or housing surrounding the plurality of fan blades ofthe fan. As such, the nacelle defines an inlet with the fuselage of theaircraft. More particularly, the inlet extends substantially around thefuselage of the aircraft when the boundary layer ingestion fan ismounted at the aft end of the aircraft. The fan assembly also includes alow-distortion inlet assembly configured with the inlet for reducingswirl distortion of the airflow entering the inlet. More specifically,the inlet assembly includes one or more inlet guide vanes configuredwithin the inlet and one or more structural members mounted atpredetermined radial locations around a circumference of the fan shaftof the fan at the inlet. Accordingly, the structural members andoptionally the inlet guide vanes are configured to reduce the swirldistortion entering the inlet of the fan.

In another aspect, the present disclosure is directed to alow-distortion inlet assembly for an aft fan mounted to a fuselage of anaircraft. The inlet assembly includes a plurality of structural membersmounted at predetermined locations around a circumference of the fanshaft of the fan at the inlet. The predetermined location(s) may bedetermined as a function of swirl distortion entering the inlet.Further, the inlet assembly may be absent of inlet guide vanes. In suchembodiments, the structural members guide the airflow axially to preventgeneration of stream-wise vorticity, thereby eliminating the need forinlet guide vanes in the inlet assembly or reducing the number of inletguide vanes needed.

In yet another aspect, the present disclosure is directed to a methodfor retrofitting a boundary layer ingestion fan for an aft end of afuselage of an aircraft with an inlet assembly configured to reduceswirl distortion at an inlet of the fan. The method includes determiningone or more locations around a circumference of a fan shaft of the fanhaving a swirl distortion exceeding a predetermined threshold. Morespecifically, the step of determining the location(s) having a swirldistortion exceeding the predetermined threshold may include usingcomputer modeling to determine which locations around the circumferenceof the inlet experience the highest swirl distortion. As such, themethod also includes replacing one or more inlet guide vanes at thepredetermined locations with one or more structural members.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 illustrates a top view of one embodiment of an aircraft accordingto the present disclosure;

FIG. 2 illustrates a port side view of the aircraft of FIG. 1;

FIG. 3 illustrates a schematic, cross-sectional view of one embodimentof a gas turbine engine mounted to one of the wings of the aircraft ofFIG. 1;

FIG. 4 illustrates a schematic, cross-sectional view of one embodimentof an aft engine according to the present disclosure;

FIG. 5 illustrates a cross-sectional view of one embodiment of an inletguide vane of an inlet assembly according to the present disclosure,particularly illustrating an inlet guide vane having a rotatable flap ata trailing edge thereof;

FIG. 6 illustrates a schematic, cross-sectional view of an aft engine ofan aircraft, viewed along an axial centerline thereof so as toillustrate one embodiment of an inlet assembly according to the presentdisclosure;

FIG. 7 illustrates a schematic, cross-sectional view of an aft engine ofan aircraft, viewed along an axial centerline thereof so as toillustrate another embodiment of an inlet assembly according to thepresent disclosure;

FIG. 8 illustrates a side view of one embodiment of an aft enginemounted to an aft end of an aircraft according to the presentdisclosure, particularly illustrating an inlet assembly having aplurality of struts integrated with the nacelle of the aft fan of theaft engine;

FIG. 9 illustrates a side view of one embodiment of an aft enginemounted to an aft end of an aircraft according to the presentdisclosure, particularly illustrating an inlet assembly having aplurality of strakes integrated with the fuselage of the aft fan of theaft engine and spaced apart from the nacelle;

FIG. 10 illustrates a schematic, cross-sectional view of an aft engineof an aircraft, viewed along an axial centerline thereof so as toillustrate another embodiment of an inlet assembly according to thepresent disclosure;

FIG. 11 illustrates a schematic, cross-sectional view of an aft engineof an aircraft, viewed along an axial centerline thereof so as toillustrate yet another embodiment of an inlet assembly according to thepresent disclosure;

FIG. 12 illustrates a cross-sectional view of another embodiment of aninlet guide vane of an inlet guide vane assembly according to thepresent disclosure, particularly illustrating an upright inlet guidevane;

FIG. 13 illustrates a cross-sectional view of yet another embodiment ofan inlet guide vane of an inlet guide vane assembly according to thepresent disclosure, particularly illustrating an inverted inlet guidevane;

FIG. 14 illustrates a cross-sectional view of still another embodimentof an inlet guide vane of an inlet guide vane assembly according to thepresent disclosure, particularly illustrating a symmetrical inlet guidevane; and

FIG. 15 illustrates a flow diagram of one embodiment of a method forretrofitting an aft fan of an aft engine with an inlet assembly so as toreduce swirl distortion at the inlet of the fan according to the presentdisclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Generally, the present disclosure is directed to a low-distortion inletassembly for reducing airflow swirl distortion entering an aft fanmounted to a fuselage of an aircraft. More specifically, the inletassembly includes a plurality of structural members, such as struts orstrakes, mounted at predetermined locations around a circumference ofthe fan shaft of the fan at the inlet. The predetermined location(s) maybe determined as a function of swirl distortion entering the inlet, e.g.using computer modeling. As such, placement of the structural member(s)at the predetermined locations is configured to reduce swirl distortionof the airflow entering the fan. In some embodiments, the inlet assemblymay also include inlet guide vanes, with some of the inlet guide vanesbeing replaced with a fewer number of the structural members so as toreduce drag and/or weight of the inlet assembly. In alternativeembodiments, the inlet assembly may be absent of inlet guide vanesaltogether.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 illustrates a top view ofone embodiment of the aircraft 10 according to the present disclosure.FIG. 2 illustrates a port side view of the aircraft 10 as illustrated inFIG. 1. As shown in FIGS. 1 and 2 collectively, the aircraft 10 definesa longitudinal centerline 14 that extends therethrough, a verticaldirection V, a lateral direction L, a forward end 16, and an aft end 18.

Moreover, the aircraft 10 includes a fuselage 12, extendinglongitudinally from the forward end 16 of the aircraft 10 towards theaft end 18 of the aircraft 10, and a pair of wings 20. As used herein,the term “fuselage” generally includes all of the body of the aircraft10, such as an empennage of the aircraft 10 and an outer surface or skin38 of the aircraft 10. The first of such wings 20 extends laterallyoutwardly with respect to the longitudinal centerline 14 from a portside 22 of the fuselage 12 and the second of such wings 20 extendslaterally outwardly with respect to the longitudinal centerline 14 froma starboard side 24 of the fuselage 12. Further, as shown in theillustrated embodiment, each of the wings 20 depicted includes one ormore leading edge flaps 26 and one or more trailing edge flaps 28. Theaircraft 10 may also include a vertical stabilizer 30 having a rudderflap 32 for yaw control, and a pair of horizontal stabilizers 34, eachhaving an elevator flap 36 for pitch control. It should be appreciatedhowever, that in other exemplary embodiments of the present disclosure,the aircraft 10 may additionally or alternatively include any othersuitable configuration of stabilizer that may or may not extend directlyalong the vertical direction V or horizontal/lateral direction L.

In addition, the aircraft 10 of FIGS. 1 and 2 includes a propulsionsystem 100, herein referred to as “system 100.” The system 100 includesa pair of aircraft engines, at least one of which mounted to each of thepair of wings 20, and an aft engine. For example, as shown, the aircraftengines are configured as turbofan jet engines 102, 104 suspendedbeneath the wings 20 in an under-wing configuration. Additionally, theaft engine is configured as an engine that ingests and consumes airforming a boundary layer over the fuselage 12 of the aircraft 10.Specifically, the aft engine is configured to ingest and consume airforming a boundary layer over the fuselage 12 of the aircraft 10, i.e.,the aft engine is a Boundary Layer Ingestion (BLI) fan 106. Further, asshown in FIG. 2, the BLI fan 106 is mounted to the aircraft 10 at alocation aft of the wings 20 and/or the jet engines 102, 104, such thata central axis 15 extends therethrough. As used herein, the “centralaxis” refers to a midpoint line extending along a length of the BLI fan106. Further, for the illustrated embodiment, the BLI fan 106 is fixedlyconnected to the fuselage 12 at the aft end 18, such that the BLI fan106 is incorporated into or blended with a tail section at the aft end18. However, it should be appreciated that in various other embodiments,some of which will be discussed below, the BLI fan 106 may alternativelybe positioned at any suitable location of the aft end 18.

In various embodiments, the jet engines 102, 104 may be configured toprovide power to an electric generator 108 and/or an energy storagedevice 110. For example, one or both of the jet engines 102, 104 may beconfigured to provide mechanical power from a rotating shaft (such as anLP shaft or HP shaft) to the electric generator 108. Additionally, theelectric generator 108 may be configured to convert the mechanical powerto electrical power and provide such electrical power to one or moreenergy storage devices 110 and/or the BLI fan 106. Accordingly, in suchembodiments, the propulsion system 100 may be referred to as agas-electric propulsion system. It should be appreciated, however, thatthe aircraft 10 and propulsion system 100 depicted in FIGS. 1 and 2 isprovided by way of example only and that in other exemplary embodimentsof the present disclosure, any other suitable aircraft 10 may beprovided having a propulsion system 100 configured in any other suitablemanner.

Referring now to FIG. 3, in certain embodiments, the jet engines 102,104 may be configured as high-bypass turbofan jet engines. Morespecifically, FIG. 3 illustrates a schematic cross-sectional view of oneembodiment of a high-bypass turbofan jet engine 200, herein referred toas “turbofan 200.” In various embodiments, the turbofan 200 may berepresentative of jet engines 102, 104. Further, as shown, the turbofan200 engine 10 defines an axial direction A₁ (extending parallel to alongitudinal centerline 201 provided for reference) and a radialdirection R₁. In general, the turbofan 200 includes a fan section 202and a core turbine engine 204 disposed downstream from the fan section202.

In particular embodiments, the core turbine engine 204 generallyincludes a substantially tubular outer casing 206 that defines anannular inlet 208. It should be appreciated, that as used herein, termsof approximation, such as “approximately,” “generally,” “substantially,”or “about,” refer to being within a thirty percent margin of error. Theouter casing 206 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 210 and ahigh pressure (HP) compressor 212; a combustion section 214; a turbinesection including a high pressure (HP) turbine 216 and a low pressure(LP) turbine 218; and a jet exhaust nozzle section 220. A high pressure(HP) shaft or spool 222 drivingly connects the HP turbine 216 to the HPcompressor 212. A low pressure (LP) shaft or spool 224 drivinglyconnects the LP turbine 218 to the LP compressor 210.

Further, as shown, the fan section 202 may include a variable pitch fan226 having a plurality of fan blades 228 coupled to a disk 230 in aspaced apart manner. As depicted, the fan blades 228 extend outwardlyfrom the disk 230 generally along the radial direction R₁. Each fanblade 228 is rotatable relative to the disk 230 about a pitch axis byvirtue of the fan blades 228 being operatively coupled to a suitableactuation member 232 configured to collectively vary the pitch of thefan blades 228 in unison. As such, the fan blades 228, the disk 230, andthe actuation member 232 are together rotatable about the longitudinalaxis 12 by LP shaft 224 across a power gearbox 234. In certainembodiments, the power gearbox 234 includes a plurality of gears forstepping down the rotational speed of the LP shaft 224 to a moreefficient rotational fan speed.

Referring still to FIG. 3, the disk 230 is covered by rotatable fronthub 236 aerodynamically contoured to promote an airflow through theplurality of fan blades 228. Additionally, the fan section 202 includesan annular fan casing or outer nacelle 238 that circumferentiallysurrounds the fan 226 and/or at least a portion of the core turbineengine 204. It should be appreciated that the outer nacelle 238 may beconfigured to be supported relative to the core turbine engine 204 by aplurality of circumferentially-spaced outlet guide vanes 240. Moreover,a downstream section 242 of the nacelle 238 may extend over an outerportion of the core turbine engine 204 so as to define a bypass airflowpassage 244 therebetween.

In addition, it should be appreciated that the turbofan engine 200depicted in FIG. 3 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 200 may have any othersuitable configuration. Further, it should be appreciated, that in otherexemplary embodiments, the jet engines 102, 104 may instead beconfigured as any other suitable aeronautical engine.

Referring now to FIG. 4, a schematic, cross-sectional side view of anaft engine in accordance with various embodiments of the presentdisclosure is provided, such as the aft engine mounted to the aircraft10 at the tail section 18 of the aircraft 10. More specifically, asshown, the aft engine is configured as a boundary layer ingestion (BLI)fan 300. The BLI fan 300 may be configured in substantially the samemanner as the BLI fan 106 described above with reference to FIGS. 1 and2 and the aircraft 10 may be configured in substantially the same manneras the exemplary aircraft 10 described above with reference to FIGS. 1and 2.

More specifically, as shown, the BLI fan 300 defines an axial directionA₂ extending along the central axis 15 that extends therethrough forreference. Additionally, the BLI fan 300 defines a radial direction R₂and a circumferential direction C₂ (FIGS. 6 and 7). In general, the BLIfan 300 includes a fan 304 rotatable about the central axis 15, anacelle 306 extending around at least a portion of the fan 304, andoptionally one or more inlet guide vanes 308 extending between thenacelle 306 and the fuselage 12 of the aircraft 10. In alternativeembodiments, the BLI fan 300 may be absent of the inlet guide vanes 308as discussed in more detail below in reference to FIG. 6. Further, thefan 304 includes a plurality of fan blades 310 spaced generally alongthe circumferential direction C₂. Moreover, where present, the inletguide vanes 308 extend between the nacelle 306 and the fuselage 12 ofthe aircraft 10 at a location forward of the plurality of fan blades310. More specifically, as shown, the inlet guide vanes 308 generallyextend substantially along the radial direction R₂ of the BLI fan 300between the nacelle 306 and the fuselage 12 of the aircraft 10 formounting the BLI fan 300 to the fuselage 12 of the aircraft 10. Inaddition, as shown in FIG. 7, the inlet guide vanes 308 may be placed ingroups 318 at predetermined locations around the circumference of thefan shaft 312 (i.e. spaced along the circumferential direction C₂ of theBLI fan 300). For example, as shown in the illustrated embodiment, theinlet assembly 302 includes two groups 318 of two inlet guide vanes 308.In alternative embodiments, the inlet assembly 302 may include more thantwo groups 318 of inlet guide vanes 308 at any circumferential locationhaving any suitable number of inlet guide vanes 308. In addition, forembodiments having more than one inlet guide vane group 318, each group318 may include the same number of inlet guide vanes 308 or a differentnumber of inlet guide vanes 308.

Further, the inlet guide vanes 308 may be shaped and/or oriented todirect and/or condition a flow of air into the BLI fan 300 to, e.g.,increase an efficiency of the BLI fan 300, or reduce a distortion of theair flowing into the BLI fan 300. In addition, it should be understoodthat the inlet guide vanes 308 may be configured as fixed inlet guidevanes extending between the nacelle 306 and the fuselage 12 of theaircraft 10. Alternatively, the inlet guide vanes 308 may be configuredas variable inlet guide vanes.

For example, as shown in FIG. 5, a cross-sectional view of one of theinlet guide vanes 308 taken along the radial direction R₂ isillustrated. As shown, the inlet guide vane 308 extends between aforward, upstream end 320 and an aft, downstream end 322. A body 325 ofthe inlet guide vane 308 is fixed relative to the nacelle 306 of the BLIfan 300 and the fuselage 12 of the aircraft 10. In addition, as shown,the inlet guide vanes 308 may also include an optional flap 324 at theaft end 320 configured to rotate about a substantially radial axis 326.For example, as shown, the flap 324 is configured to rotate between afirst position 328 (in phantom), a neutral position 330, a secondposition 332 (in phantom), and a potentially infinite number ofpositions therebetween. By rotating the flap 324 between the variouspositions, the inlet guide vanes 308 may be configured to vary adirection in which air flowing thereover is directed.

Additionally, the nacelle 306 extends around and encircles the pluralityof fan blades 310, and also extends around the fuselage 12 of theaircraft 10 at an aft end 18 of the aircraft 10 when, as shown in FIG.4, the BLI fan 300 is mounted to the aircraft 10. Notably, as usedherein, the term “nacelle” includes the nacelle as well as anystructural fan casing or housing. As is also depicted in FIG. 4, the fan304 additionally includes a fan shaft 312 with the plurality of fanblades 310 attached thereto. Although not depicted, the fan shaft 312may be rotatably supported by one or more bearings located forward ofthe plurality of fan blades 310 and, optionally, one or more bearingslocated aft of the plurality of fan blades 310. Such bearings may be anysuitable combination of roller bearings, ball bearings, thrust bearings,etc.

In certain embodiments, the plurality of fan blades 310 may be attachedin a fixed manner to the fan shaft 312, or alternatively, the pluralityof fan blades 310 may be rotatably attached to the fan shaft 312. Forexample, the plurality of fan blades 310 may be attached to the fanshaft 312 such that a pitch of each of the plurality of fan blades 310may be changed, e.g., in unison, by a pitch change mechanism (notshown). Changing the pitch of the plurality of fan blades 310 mayincrease an efficiency of the BLI fan 300 and/or may allow the BLI fan300 to achieve a desired thrust profile. With such an embodiment, theBLI fan 300 may be referred to as a variable pitch BLI fan.

The fan shaft 312 is mechanically coupled to a power source 314 locatedat least partially within the fuselage 12 of the aircraft 10, forward ofthe plurality of fan blades 310. Further, as shown, the fan shaft 312 ismechanically coupled to the power source 314 through a gearbox 316. Thegearbox 316 may be configured to modify a rotational speed of the powersource 314, or rather of a shaft 315 of the power source 314, such thatthe fan 304 of the BLI fan 300 rotates at a desired rotational speed.The gearbox 316 may be a fixed ratio gearbox, or alternatively, thegearbox 316 may define a variable gear ratio. With such an embodiment,the gearbox 316 may be operably connected to, e.g., a controller of theaircraft 10 for changing its ratio in response to one or more flightconditions.

In certain embodiments, the BLI fan 300 may be configured with agas-electric propulsion system, such as the gas-electric propulsionsystem 100 described above with reference to FIG. 1. In such anembodiment, the power source 314 may be an electric motor that receivespower from one or both of an energy storage device or an electricgenerator—such as the energy storage device 110 or electric generator108 of FIGS. 1 and 2, the electric generator 108 converting mechanicalpower received from one or more under-wing mounted aircraft engines toelectric power. However, in other embodiments, the power source 314 mayinstead be any other suitable power source. For example, the powersource 314 may alternatively be configured as a gas engine, such as agas turbine engine or internal combustion engine. Moreover, in certainexemplary embodiments, the power source 314 may be positioned at anyother suitable location within, e.g., the fuselage 12 of the aircraft 10or the BLI fan 300. For example, in certain embodiments, the powersource 314 may be configured as a gas turbine engine positioned at leastpartially within the BLI fan 300.

Referring still to FIG. 4, the BLI fan 300 may also additionally includeone or more outlet guide vanes 338 and a tail cone 340. As shown in theillustrated embodiment, the outlet guide vanes 338 extend between thenacelle 306 and the tail cone 340 for directing a flow of air throughthe BLI fan 300, and optionally for adding strength and rigidity to theBLI fan 300. Further, the outlet guide vanes 338 may be evenly spacedalong the circumferential direction C₂ or may have any other suitablespacing. Additionally, the outlet guide vanes 338 may be fixed outletguide vanes, or alternatively may be variable outlet guide vanes.

Further, aft of the plurality of fan blades 310, and for the embodimentdepicted, aft of the one or more outlet guide vanes 338, the BLI fan 300additionally defines a nozzle 342 between the nacelle 306 and the tailcone 340. As such, the nozzle 342 may be configured to generate anamount of thrust from the air flowing therethrough. In addition, thetail cone 340 may be shaped to minimize an amount of drag on the BLI fan300. However, in other embodiments, the tail cone 340 may have any othershape and may, e.g., end forward of an aft end of the nacelle 306 suchthat the tail cone 340 is enclosed by the nacelle 306 at an aft end.Additionally, in other embodiments, the BLI fan 300 may not beconfigured to generate any measurable amount of thrust, and instead maybe configured to ingest air from a boundary layer of air of the fuselage12 of the aircraft 10 and add energy/speed up such air to reduce anoverall drag on the aircraft 10 (and thus increase a net thrust of theaircraft 10).

Referring particularly to FIGS. 4 and 6-7, the BLI fan 300 defines aninlet 334 at a forward end 336 between the nacelle 306 and the fuselage12 of the aircraft 10. As mentioned above, the nacelle 306 of the BLIfan 300 extends around the central axis 15 of the aircraft 10 and thefuselage 12 of the aircraft 10 at the aft end of the aircraft 10.Accordingly, as shown, the inlet 334 of the BLI fan 300 extendssubstantially three hundred sixty degrees (360°) around the central axis15 of the aircraft 10 and the fuselage 12 of the aircraft 10 when, suchas in the embodiment depicted, the BLI fan 300 is mounted to theaircraft 10. Additionally, in still further embodiments, the BLI fan300, or rather the external surface of the nacelle 306, may have anyother suitable cross-sectional shapes along the axial direction A₂ (asopposed to the circular shape depicted) and the structural members 307may not be evenly spaced along the circumferential direction C₂.

Referring particularly to FIGS. 6 and 7, schematic, cross-sectionalviews of various embodiments of the BLI fan 300, viewed along an axialcenterline 15 thereof so as to illustrate a low-distortion inletassembly 302 configured with the inlet 334 of the fan 300 according tothe present disclosure are illustrated. More specifically, as shown, theinlet assembly 302 includes a plurality of structural members 307mounted at one or more predetermined radial locations around acircumference of the fan shaft 312 of the BLI fan 300. For example, asshown in FIG. 10, the predetermined radial locations as described hereinmay have a swirl distortion exceeding a predetermined threshold asillustrated by arrows 352. Further, it should be understood that thepredetermined radial locations may be at the illustrated locations aswell as any location therebetween and are meant to encompass locationshaving a high swirl distortion and/or a location where a modification ofthe airflow would have the highest impact of correcting the swirldistortion. In other words, for certain embodiments, the airflowentering the BLI fan 300 may be evaluated to determine a swirl patternthereof. If the swirl distortion is deemed to cause damage to the fan300, the structural members 307 can be chosen so as to reduce the swirldistortion and reduce such damage. Thus, the location and/or number ofstructural members 307, as well as the shape and orientation of thestructural members 307, may be designed and chosen as a function of theswirl pattern. Accordingly, the structural member(s) 307 are configuredto condition the airflow into the BLI fan 300 to, e.g., increase anefficiency of the BLI fan 300, or reduce a distortion of the air flowinginto the BLI fan 300, which will be discussed in more detail below.

More specifically, as shown in FIGS. 6-10, the structural member(s) 307may be configured as struts 309 (FIGS. 6-8 and 10), strakes 311 (FIG.9), or any other suitable structural component. Generally, struts arestructural components designed to resist longitudinal compression. Inaddition, the struts 309 of the present disclosure are configured toguide the airflow entering the BLI fan 300 axially, thereby preventinggeneration of stream-wise vorticity and eliminating the need of inletguide vanes 308 in certain embodiments. Further, strakes generally referto wing-like projections, airfoils, or vanes used to improve aerodynamicefficiency. As such, the struts 309 and/or strakes 311 of the presentdisclosure are strategically placed at the predetermined locations so asto redistribute the airflow entering the fan 300 more uniformlycircumferentially so as to reduce swirl distortion at the inlet 334. Inparticular embodiments, the structural members 307 may be non-uniformlyspaced along the circumferential direction C₂ of the BLI fan 300 aroundthe fan shaft 312. In alternative embodiments, the structural members307 may be evenly spaced along the circumferential direction C₂ of theBLI fan 300 around the fan shaft 312.

For example, as shown in FIG. 6, the illustrated inlet assembly 302includes at least four struts 309 non-uniformly spaced around thecircumference of the fan shaft 312. In further embodiments, the inletassembly 302 may include more than four or less than four struts 309. Inaddition, as shown in FIG. 7, the inlet assembly 302 may include acombination of struts 309 and inlet guide vanes 308 spaced around thecircumference of the fan shaft 312. In yet another embodiment, as shownin FIG. 9, the inlet assembly 302 may include a plurality of strakes311.

In additional embodiments, the structural member(s) 307 may beintegrated with at least one of the nacelle 306 of the fan 300 or thefuselage 12. More specifically, as shown in FIG. 8, the structuralmember(s) 307 may be integrated with the nacelle 306 so as to provideadditional strength and/or support thereto. In addition, as shown inFIG. 11, the structural member(s) 307 may be integrated with and extendpast the nacelle 306 of the BLI fan 300 in the radial direction R₂, e.g.as shown via struts 309. In such embodiments, the structural member(s)307 may be configured to provide an aircraft control surface as well asreduce distortion entering the BLI fan 300. In further embodiments, asshown in FIG. 8, the structural member(s) 307 may extend in front of thenacelle 306 of the fan 300 in the axial direction A₂.

Alternatively, as shown in FIG. 9, the structural member(s) 307, e.g.strakes 311, may be integrated with the fuselage 12. For example, asshown in the illustrated embodiment, the illustrated strakes 311 arespaced apart from the nacelle 306 by a distance D and are locatedupstream of the nacelle 306 of the BLI fan 300 in the axial directionA₂.

Referring now to FIGS. 12-14, cross-sectional views of variousembodiments of the structural member(s) 307 of the present disclosureare illustrated. As shown, each of the structural member(s) 307 may havea unique shape and/or orientation corresponding to airflow conditionsentering the BLI fan 300 at a particular location in the fan 300. Assuch, the structural member(s) 307 may be cambered to mitigate the swirldistortion and/or to pre-swirl the flow entering the fan 300. Thus, anycombination of shapes may be used in the inlet assembly 302 and can bechosen based on a determined swirl distortion of the airflow enteringthe BLI fan 300.

It should be understood that the lift force generated by the structuralmember(s) 307 depends on the shape of its cross-section, especially theamount of camber (i.e. curvature such that the upper surface is moreconvex than the lower surface). In other words, increasing the cambergenerally increases lift. Therefore, the structural member(s) 307 can betailored to reduce fan flow distortion by introducing variations of themember(s) 307. For example, as shown in FIGS. 11 and 12, one or more ofthe structural member(s) 307 may include a cambered cross-sectionalshape configured to reduce swirl distortion entering the BLI fan 300.More specifically, as shown, the structural member(s) 307 may have acambered upright airfoil cross-section (FIG. 12), a cambered invertedairfoil cross-section (FIG. 13), or a symmetrical airfoil cross-section(FIG. 14).

More specifically, as shown in FIG. 12, the cambered upright structuralmember 307 generally has a mean camber line 348 above the chord line 350of the airfoil, with the trailing edge 346 of the structural member 307having a downward direction. Such cambered airfoils typically generatelift at zero angle of attack and since air follows the trailing edge346, the air is deflected downward. As shown in FIG. 13, the invertedstructural member 307 generally has a mean camber line 348 below thechord line 350 of the airfoil, with the trailing edge 346 of thestructural member 307 having an upward direction. When a camberedairfoil is upside down, the angle of attack can be adjusted so that thelift force is upwards. In contrast, as shown in FIG. 14, the mean camberline 348 and the chord line 350 of a symmetrical structural member 307are the same (i.e. the lines 348, 350 overlap and there is zero camber).

Referring now to FIG. 15, a flow diagram of one embodiment of a method400 for retrofitting the BLI fan 300 with the inlet assembly 302 of thepresent disclosure so as to reduce swirl distortion at the inlet 334 ofthe BLI fan 300 is illustrated. As shown at 402, the method 400 includesdetermining one or more locations around a circumference of the fanshaft 312 of the BLI fan 300 having a swirl distortion exceeding apredetermined threshold. More specifically, the step of determining thelocation(s) having a swirl distortion exceeding the predeterminedthreshold may include using computer modeling and/or system monitoringto determine which locations around the circumference of the inletexperience the highest swirl distortion. As shown at 404, the method 400includes replacing one or more inlet guide vanes 308 at thepredetermined locations with one or more structural members 307. Morespecifically, as mentioned, the structural member(s) 307 may includestruts 309, strakes 311, or any other suitable structural component.Further, in one embodiments, the method 400 may include replacing all ofthe inlet guide vanes with a fewer number of structural members 307. Assuch, the retrofitted inlet assembly 302 is configured to reduce dragand/or weight of the aircraft 10, while also reducing swirl distortionentering the BLI fan 300.

In yet another embodiment, it should be understood that the presentdisclosure may also include a method for initially assembling the BLIfan 300 with an inlet assembly 302 that reduces swirl distortion at theinlet 334 of the BLI fan 300. For example, in such an embodiment, themethod may include determining one or more locations around acircumference of the fan shaft 312 of the BLI fan 300 having a swirldistortion exceeding a predetermined threshold. Further, the methodincludes mounting one or more structural members 307 and optionally oneor more inlet guide vanes 308 at the predetermined locations. As such,the structural members 307 can be used alone or in combination with afewer number of inlet guide vanes 308 so as to reduce drag and/or weightof the aircraft 10, while also reducing swirl distortion entering theBLI fan 300.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A boundary layer ingestion fan assembly formounting to an aft end of a fuselage of an aircraft, the boundary layeringestion fan assembly comprising: a fan rotatable about a central axisof the boundary layer ingestion fan, the fan comprising a plurality offan blades rotatable about a fan shaft; a nacelle surrounding theplurality of fan blades of the fan, the nacelle defining an inlet withthe fuselage of the aircraft, the inlet extending substantially aroundthe fuselage of the aircraft when the boundary layer ingestion fan ismounted at the aft end of the aircraft; a low-distortion inlet assemblyconfigured with the inlet, the inlet assembly comprising: one or moreinlet guide vanes configured within the inlet; and one or morestructural members mounted at predetermined radial locations around acircumference of the fan shaft of the fan at the inlet, the structuralmembers configured to reduce a swirl distortion entering the inlet ofthe fan.
 2. The boundary layer ingestion fan assembly of claim 1,wherein the one or more structural members comprise at least one of astrut or a strake.
 3. The boundary layer ingestion fan assembly of claim1, wherein one or more of the structural members are integrated with atleast one of the nacelle of the fan or the fuselage.
 4. The boundarylayer ingestion fan assembly of claim 3, wherein the one or morestructural members are integrated with and extend past the nacelle ofthe fan in a radial direction.
 5. The boundary layer ingestion fanassembly of claim 1, wherein one or more of the structural membersextend in front of the nacelle of the fan in an axial direction.
 6. Theboundary layer ingestion fan assembly of claim 1, wherein one or more ofthe structural members are spaced apart from and upstream of the nacelleof the fan in an axial direction.
 7. The boundary layer ingestion fanassembly of claim 1, wherein one or more of the structural memberscomprises a cambered cross-sectional shape configured to reduce swirldistortion entering the fan.
 8. The boundary layer ingestion fanassembly of claim 1, wherein one or more of the structural members arenon-uniformly spaced around the circumference of the fan shaft.
 9. Theboundary layer ingestion fan assembly of claim 1, further comprising aplurality of inlet guide vanes placed in groups around the circumferenceof the fan shaft.
 10. A low-distortion inlet assembly for an aft fanmounted to a fuselage of an aircraft, the inlet assembly comprising: aplurality of structural members mounted at predetermined radiallocations around a circumference of the fan shaft of the fan at theinlet, the structural members configured to reduce a swirl distortionentering the inlet of the fan; and wherein the inlet assembly is absentof inlet guide vanes.
 11. The inlet assembly of claim 10, wherein theone or more structural members comprise at least one of a strut or astrake.
 12. The inlet assembly of claim 10, wherein one or more of thestructural members are integrated with at least one of the nacelle ofthe fan or the fuselage.
 13. The inlet assembly of claim 12, wherein theone or more structural members are integrated with and extend past thenacelle of the fan in a radial direction.
 14. The inlet assembly ofclaim 10, wherein one or more of the structural members extend in frontof the nacelle of the fan in an axial direction.
 15. The inlet assemblyof claim 10, wherein one or more of the structural members are spacedapart from and located upstream of the nacelle of the fan in an axialdirection.
 16. The inlet assembly of claim 10, wherein one or more ofthe structural members comprises a cambered cross-sectional shapeconfigured to reduce swirl distortion entering the fan.
 17. The inletassembly of claim 10, wherein one or more of the structural members arenon-uniformly spaced around the circumference of the fan shaft.
 18. Amethod for retrofitting a boundary layer ingestion fan for an aft end ofa fuselage of an aircraft with an inlet assembly configured to reduceswirl distortion at an inlet of the fan, the method comprising:determining one or more radial locations around a circumference of a fanshaft of the fan having a swirl distortion exceeding a predeterminedthreshold; and replacing one or more inlet guide vanes at thepredetermined locations with one or more structural members.
 19. Themethod of claim 18, wherein the one or more structural members compriseat least one of a strut or a strake.
 20. The method of claim 18, furthercomprising replacing all of the inlet guide vanes with a fewer number ofstructural members.